1. Field of the Invention
The present invention relates generally to gas turbine engines and, more particularly, to a full-round compressor casing assembly in a gas turbine engine.
2. Description of the Prior Art
Gas turbine engines typically include a core engine having a compressor for compressing air entering the core engine, a combustor where fuel is mixed with the compressed air and then burned to create a high energy gas stream, and a first or high pressure turbine which extracts energy from the gas stream to drive the compressor. In aircraft turbofan engines, a second turbine or low pressure turbine located downstream from the core engine extracts more energy from the gas stream for driving a forward fan. The forward fan provides the main propulsive thrust generated by the engine.
The compressor typically includes multiple alternating axially-arranged stages of movable blades and stationary vanes. Each stage of movable blades includes a row of blades attached to one of a plurality of rotating rotor discs. Each stage of stationary vanes includes a row of vanes attached to an outer casing encompassing the stages of movable blades and stationary vanes.
Outer casings of compressors typically fall generally in three different prior art design categories: a split line 180.degree. assembly, a sector assembly, and a bolted stage assembly. In the split line 180.degree. casing assembly design, the vanes are assembled into two casing halves and then joined around the rotor structure by means of two horizontal split line flanges. In the sector casing assembly design, vane sectors are assembled around the rotor structure for all stages, then a full-round casing is slipped over the assembly to lock all sectors in place. In the bolted stage casing assembly, full-round stator nozzle assemblies which have rotor shrouding cantilevered off the stator are stacked with the rotor structure and then fastened together by bolted joints.